Coriolis optimized U-channel with root flag

ABSTRACT

An airfoil includes pressure and suction side walls that extend in a chord-wise direction between a leading edge and a trailing edge. The pressure and suction side walls extend in a radial direction between a platform and a tip to provide an exterior airfoil surface. A cooling passage is arranged between the pressure and suction side walls and has a first passage along the pressure side wall and a second passage along the suction side wall. The first passage is configured to receive cooling air from a cooling air source radially inward of the platform. The second passage is configured to receive cooling air from the first passage near the tip. A root flag passage is configured to purge the cooling air from the second passage near the trailing edge.

BACKGROUND

This disclosure relates to gas turbine engines and particularly tointernally cooled rotor blades.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-temperature and pressure gas flow. The hot gas flow expands throughthe turbine section to drive the compressor and the fan section.

As is well known, the aircraft engine industry is experiencing asignificant effort to improve the gas turbine engine's performance whilesimultaneously decreasing its weight. The ultimate goal has been toattain the optimum thrust-to-weight ratio. One of the primary areas offocus to achieve this goal is the “hot section” of the engine since itis well known that engine's thrust/weight ratio is significantlyimproved by increasing the temperature of the turbine gases. However,turbine gas temperature is limited by the metal temperature constraintsof the engine's components. Significant effort has been made to achievehigher turbine operating temperatures by incorporating technologicaladvances in the internal cooling of the turbine blades.

Serpentine core cooling passages have been used to cool turbine blades.An example serpentine cooling passage is arranged between the leadingand trailing edge core cooling passages in a chord-wise direction. Sucha configuration typically provides “up” passages arranged near theleading and trailing edges fluidly joined by a “down” passage. In sucharrangements, the Coriolis effect may augment the heat transfercoefficient on the pressure side of an up pass and the suction side of adown pass. With a conventional serpentine design, this only allows onehot wall to take advantage of this augmentation.

SUMMARY

In one exemplary embodiment, an airfoil includes pressure and suctionside walls that extend in a chord-wise direction between a leading edgeand a trailing edge. The pressure and suction side walls extend in aradial direction between a platform and a tip to provide an exteriorairfoil surface. A cooling passage is arranged between the pressure andsuction side walls and has a first passage along the pressure side walland a second passage along the suction side wall. The first passage isconfigured to receive cooling air from a cooling air source radiallyinward of the platform. The second passage is configured to receivecooling air from the first passage near the tip. A root flag passage isconfigured to purge the cooling air from the second passage near thetrailing edge.

In a further embodiment of the above, the root flag passage is arrangedalong the suction side wall.

In a further embodiment of any of the above, the root flag passage isarranged radially outward from the platform.

In a further embodiment of any of the above, the root flag passageextends more than 50% of a chord length.

In a further embodiment of any of the above, the first and secondcooling passages have the same cross sectional shape.

In a further embodiment of any of the above, a cross section of thecooling passage is substantially rectangular.

In a further embodiment of any of the above, a cross section of thecooling passage is substantially triangular.

In a further embodiment of any of the above, the cooling passage has anaspect ratio of less than two.

In a further embodiment of any of the above, the first and secondpassages are arranged at a same position in the chord-wise direction.

In a further embodiment of any of the above, the cooling air source isbleed air from a compressor section of a gas turbine engine.

In another exemplary embodiment, a gas turbine engine includes acombustor section that is arranged fluidly between compressor andturbine sections. An airfoil is arranged in the turbine section. Theairfoil including pressure and suction side walls that extend in achord-wise direction between a leading edge and a trailing edge. Thepressure and suction side walls extend in a radial direction between aplatform and a tip to provide an exterior airfoil surface. A coolingpassage is arranged between the pressure and suction side walls and hasa first passage along the pressure side wall and a second passage alongthe suction side wall. The first passage is configured to receivecooling air from a cooling air source. The second passage is configuredto receive cooling air from the first passage near a tip of the airfoiland a root flag passage is configured to purge the cooling air from thesecond passage near the trailing edge.

In a further embodiment of any of the above, the cooling air source isbleed air from a compressor section of a gas turbine engine.

In a further embodiment of any of the above, the root flag passage isarranged along the suction side wall.

In a further embodiment of any of the above, the root flag passage isarranged radially outward from the platform.

In a further embodiment of any of the above, the root flag passageextends more than 50% of a chord length.

In a further embodiment of any of the above, a cross section of thecooling passage is substantially rectangular.

In a further embodiment of any of the above, a cross section of thecooling passage is substantially triangular.

In a further embodiment of any of the above, the cooling passage has anaspect ratio of less than two.

In a further embodiment of any of the above, the first and secondpassages are arranged at a same position in the chord-wise direction.

In a further embodiment of any of the above, the airfoil is arranged ina first stage of the turbine section and further comprises a hybridcavity along one of the pressure and suction side walls.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 is a schematic view of an example gas turbine engine according toa first non-limiting example.

FIG. 2A is a perspective view of an airfoil.

FIG. 2B is a plan view of the airfoil illustrating directionalreferences.

FIG. 3 is a sectional view taken along line 3-3 of FIG. 2A.

FIG. 4 is a sectional view taken along line 4-4 of FIG. 2A.

FIG. 5 is a sectional view taken along line 5-5 of FIG. 2A.

FIG. 6 is another embodiment of the disclosed airfoil.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle15, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects, a first (or low) pressure compressor 44 and a first (orlow) pressure turbine 46. The inner shaft 40 is connected to the fan 42through a speed change mechanism, which in exemplary gas turbine engine20 is illustrated as a geared architecture 48 to drive a fan 42 at alower speed than the low speed spool 30. The high speed spool 32includes an outer shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high) pressure turbine 54. Acombustor 56 is arranged in exemplary gas turbine engine 20 between thehigh pressure compressor 52 and the high pressure turbine 54. Amid-turbine frame 57 of the engine static structure 36 may be arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of the low pressure compressor, or aftof the combustor section 26 or even aft of turbine section 28, and fan42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1 and less than about 5:1. Itshould be understood, however, that the above parameters are onlyexemplary of one embodiment of a geared architecture engine and that thepresent invention is applicable to other gas turbine engines includingdirect drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (′TSFC)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2A illustrates an example turbine blade 64. A root 74 of eachturbine blade 64 is mounted to a rotor disk. The turbine blade 64includes a platform 76, which provides the inner flow path, supported bythe root 74. An airfoil 78 extends in a radial direction R (shown inFIG. 2B) from the platform 76 to a tip 80. It should be understood thatthe turbine blades may be integrally formed with the rotor such that theroots are eliminated. In such a configuration, the platform is providedby the outer diameter of the rotor. The airfoil 78 provides leading andtrailing edges 82, 84. The tip 80 is arranged adjacent to a blade outerair seal (not shown).

The airfoil 78 of FIGS. 2A and 2B somewhat schematically illustrate anexterior airfoil surface 79 extending in a chord-wise direction C from aleading edge 82 to a trailing edge 84. The airfoil 78 is providedbetween pressure (typically concave) and suction (typically convex)walls 86, 88 in an airfoil thickness direction T, which is generallyperpendicular to the chord-wise direction C. Multiple turbine blades 64are arranged circumferentially in a circumferential direction A. Theairfoil 78 extends from the platform 76 in the radial direction R, orspanwise, to the tip 80.

FIG. 3 illustrates a cross-sectional view of the airfoil 78. The airfoil78 has a U-shaped cooling passage 90 arranged at a position in thechord-wise direction C. The cooling passage 90 comprises a first passage90 a arranged on the pressure side 86, and a second passage 90 barranged on the suction side 88. Throughout this disclosure, “up”passages refer to cooling passages that transport cooling fluid radiallyoutward away from the engine centerline, in a direction towards a largerradial outboard location. Conversely, “down” passages, refer to coolingpassages that transport cooling fluid radially inward toward the enginecenterline, in a direction towards a smaller inboard location. The firstpassage 90 a is an up passage, and is fluidly connected to the secondpassage 90 b, which is a down passage, near the tip 80 at a passage 90c. The first passage 90 a receives cooling air from a cooling source 92,such as bleed air from the compressor section 24. The first and secondpassages 90 a, 90 b are separated by a rib 91.

The Coriolis effect acts on cooling fluid as it is in motion relative toa rotating component, such as an airfoil. Here, as the turbine blade 64rotates about the engine central longitudinal axis A, inertia pushes thecooling air against a wall of the blade 64, which enhances the coolingon that wall. In typical cooling passage arrangements, the Corioliseffect augments the heat transfer coefficient on the pressure side of anup pass and the suction side of a down pass. The disclosed U-shapedcooling passage arrangement takes advantage of the Coriolis effect onboth the pressure and suction sides 86, 88 of the airfoil 78 as thecooling air moves circumferentially across the airfoil 78. The Corioliseffect improves the heat transfer coefficient on the pressure side 86 ascooling air travels up the first passage 90 a and on the suction side 88as cooling air travels down the second passage 90 b.

Referring now to FIG. 4, the airfoil 78 may include additional coolingpassages 94, 96, 98, 100, 102. In an example embodiment, coolingpassages 94, 98 are “up” passages, and cooling passages 100, 102 are“down” passages. The location of the U-shaped passage 90 within theairfoil 78 is selected such that the thickness of the airfoil 78 canaccommodate both the first (“up”) passage 90 a and the second (“down”)passage 90 b at the same position in the chord-wise direction C. Theillustrated example has a single U-shaped passage 90 within the airfoil.In some examples, additional U-shaped passages may be provided at otherpositions in the chord-wise direction C.

In the illustrated embodiment, the first and second passages 90 a, 90 bhave a rectangular cross section. In this embodiment, the first andsecond passages 90 a, 90 b have the same cross-sectional shape. Thefirst and second passages 90 a, 90 b have a width W oriented generallyalong the chord-wise direction C, and a height H oriented generallyalong the thickness direction T. In an embodiment, an aspect ratio ofthe width W to the height H is less than two.

Referring now to FIG. 5, the second passage 90 b is fluidly connected toa root flag passage 104. Cooling air travels from the cooling source 92,up the first passage 90 a to the airfoil tip 80, and down the secondpassage 90 b. The root flag passage 104 dumps the cooling air from thesecond passage 90 b at a root flag outlet 105 near the airfoil trailingedge 84. In one embodiment, the root flag outlet 105 purges the coolingair aft of the U-shaped passage 90. In a further embodiment, the rootflag outlet 105 purges the cooling air through the trailing edge 84. Theroot flag passage 104 is arranged near the platform 76, and runs alongthe suction side of the airfoil 78. The root flag passage 104 may extendat least 50% of the length of the airfoil 78. In an embodiment, the rootflag passage 104 is arranged just radially outward from the platform 76.In a further embodiment, the root flag passage 104 is arranged at aradially innermost portion of the airfoil 78. The root flag passage 104does not fluidly communicate with the passages 94, 96, 98, 100, 102.This purging of the cooling air through the root flag passage 104 intothe wake region aft of the airfoil trailing edge 84 may have a lowerefficiency loss than purging the flow as a suction side film.

FIG. 6 illustrates another embodiment of the airfoil 78. In thisexample, the first and second passages 90 a, 90 b have a triangularcross section. The exterior airfoil surface 79 may include film coolingholes in fluid communication with the cooling passage 90 to create athin film boundary layer that protects the exterior airfoil 78 from hotgases in the core flow path C. In some embodiments, the airfoil 78 mayalso include at least one hybrid cavity passage 106. In the illustratedembodiment, the hybrid cavity passage 106 is on the pressure side 86.Hybrid cavity passages 106 may be oriented along the pressure and/orsuction sides 86, 88, and may help shield the root flag passage 104 andreduce heat pick up. Hybrid cavity passage 106 extends radially and isprovided in a thickness direction T between the cooling passage 90 andthe airfoil surface 79 on at least one of the pressure and suction sides86, 88. In some embodiments, additional hybrid cavity passages may beprovided forward or aft of the cooling passage 90. Hybrid cavitypassages 106 have a much higher width to height aspect ratio thanpassage 90. Hybrid cavity passages 106 may protect the cooling passage90 and root flag passage 104 from gaining heat from the core flow pathC. Hybrid cavities may be particularly helpful for first stage blades,for example.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiments, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the embodiments of the presentinvention. Additionally it is important to note that any complexmulti-facetted resupply geometries that bridge centrally located mainbody cooling passages and peripherally located hybrid skin core coolingcavity passages can be created at any number of radial, circumferential,and/or tangential locations within an internal cooling configuration.The quantity, size, orientation, and location will be dictated by thenecessity to increase the local thermal cooling effectiveness andachieve the necessary thermal performance required to mitigate hotsection part cooling airflow requirements, as well as, meet part andmodule level durability life, stage efficiency, module, and overallengine cycle performance and mission weight fuel burn requirements.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil comprising: pressure and suction sidewalls extending in a chord-wise direction between a leading edge and atrailing edge, the pressure and suction side walls extending in a radialdirection between a platform and a tip to provide an exterior airfoilsurface; a cooling passage arranged between the pressure and suctionside walls having a first passage along the pressure side wall and asecond passage along the suction side wall, the first and secondpassages separated by a rib; the first passage configured to receivecooling air from a cooling air source radially inward of the platform,and the second passage configured to receive the cooling air from thefirst passage adjacent the tip; and a root flag passage configured topurge the cooling air from the second passage through the trailing edge.2. The airfoil according to claim 1, wherein the root flag passage isarranged along the suction side wall.
 3. The airfoil according to claim1, wherein the root flag passage is arranged radially outward from theplatform.
 4. The airfoil according to claim 1, wherein the root flagpassage extends more than 50% of a chord length.
 5. The airfoilaccording to claim 1, wherein the first and second passages have a samecross sectional shape.
 6. The airfoil according to claim 1, wherein across section of the cooling passage is substantially rectangular. 7.The airfoil according to claim 1, wherein a cross section of the coolingpassage is substantially triangular.
 8. The airfoil according to claim1, wherein the cooling passage has an aspect ratio of less than two. 9.The airfoil according to claim 1, wherein the first and second passagesare arranged at a same position in the chord-wise direction.
 10. Theairfoil according to claim 1, wherein the cooling air source is bleedair from a compressor section of a gas turbine engine.
 11. A gas turbineengine comprising: a combustor section arranged fluidly betweencompressor and turbine sections; and an airfoil arranged in the turbinesection, the airfoil including pressure and suction side walls extendingin a chord-wise direction between a leading edge and a trailing edge,the pressure and suction side walls extending in a radial directionbetween a platform and a tip to provide an exterior airfoil surface, acooling passage arranged between the pressure and suction side wallshaving a first passage along the pressure side wall and a second passagealong the suction side wall, the first and second passages separated bya rib, wherein the first passage is configured to receive cooling airfrom a cooling air source, and the second passage is configured toreceive the cooling air from the first passage adjacent the tip of theairfoil and a root flag passage configured to purge the cooling air fromthe second passage through the trailing edge.
 12. The gas turbine engineaccording to claim 11, wherein the cooling air source is bleed air fromthe compressor section of the gas turbine engine.
 13. The gas turbineengine according to claim 11, wherein the root flag passage is arrangedalong the suction side wall.
 14. The gas turbine engine according toclaim 11, wherein the root flag passage is arranged radially outwardfrom the platform.
 15. The gas turbine engine according to claim 11,wherein the root flag passage extends more than 50% of a chord length.16. The gas turbine engine according to claim 11, wherein a crosssection of the cooling passage is substantially rectangular.
 17. The gasturbine engine according to claim 11, wherein a cross section of thecooling passage is substantially triangular.
 18. The gas turbine engineaccording to claim 11, wherein the cooling passage has an aspect ratioof less than two.
 19. The gas turbine engine according to claim 11,wherein the first and second passages are arranged at a same position inthe chord-wise direction.
 20. The gas turbine engine according to claim11, wherein the airfoil is arranged in a first stage of the turbinesection and further comprises a hybrid cavity along one of the pressureand suction side walls.